The present invention generally relates to gas turbine engines and, more particularly, to turbine nozzles.
A gas turbine engine includes a compressor, a combustor, and a turbine. The compressor provides compressed air to the combustor. The combustor mixes the compressed air with fuel, ignites the mixture, and provides combustion gases to the turbine. The turbine extracts energy from the combustion gases.
The turbine includes one or more stages with each stage having an annular turbine nozzle and a plurality of rotor blades. The turbine nozzle channels the combustion gases to the rotor blades and the rotor blades extract energy from the combustion gases.
The turbine nozzle comprises a plurality of circumferentially spaced stator vanes positioned between and attached to radially inner and outer bands (endwalls). The circumferentially spaced vanes define converging channels there between through which the combustion gases are turned and accelerated toward the rotor blades.
Cooled turbine nozzle castings are typically one of the critical path components in gas turbine engine fabrication. In engine development programs, the first engine to test date is limited by the long schedule required to fabricate the cooled high pressure turbine (HPT) blade and nozzle parts. Due to the expensive tooling and fabrication cost of the cooled nozzle, limited quantities of hardware are purchased for development programs. A critical engine design parameter is the minimum flow area of the nozzle (nozzle area), which affects the operating efficiency of the turbine and the entire engine. After engine testing, it is often discovered that the nozzle area requirement to match the engine as a system (for optimal performance) is different than the nozzle area purchased. Thus, additional hardware must be purchased to match the engine for optimal performance.
When multiple nozzle area class sizes are purchased for program “risk mitigation”, several classes do not get utilized because they are not the needed size class at the end of the program. This is a waste of expensive hardware, tooling, and sometimes program cycle time. In some cases, the program does not have additional hardware assets (due to cost constraints) to achieve optimal engine matching of turbine nozzle areas, resulting in a specific fuel consumption increase in the “as tested” demonstrator. This can be a significant customer satisfaction issue when, for example, the specific fuel consumption increase is 2% or more.
Due to the critical nature of the nozzle area and the expense of the hardware, methods of modifying the nozzle area have been described. Many of the disclosed methods utilize some form of “airfoil rotation” to adjust the nozzle area. Known designs include rotating the entire vane, rotating just the aft portion of the vane, rotating all the vanes, and rotating just some of the vanes. These techniques have required rotational attachment apparatus and actuation mechanisms for rotating the vanes or just their aft ends. Known attachment apparatus include a shaft connected to the vane and moveably attached to an actuation mechanism. For example, U.S. Pat. No. 6,736,595 describes the use of airfoil rotation in conjunction with lever plates to modify the nozzle area. In this method, the vanes are connected to coupling shafts, which in turn are connected to link plates. The link plates are movably connected to a lever plate. Moving the link plates relative to the lever plate rotates the vanes. Although this method may be used to modify the nozzle area, performance of the engine may be decreased because the optimal vector diagram to the blade is not maintained. Performance degradation may also occur due to leakages between the airfoils and endwalls. Additionally, this method does not address the problems associated with channel variation which will occur in the airfoils, endwall, and rotation linkage mechanisms. Furthermore, this method of adjusting airflow is not viable for integral airfoil and endwall assemblies when superalloys airfoils are coated with thermal barrier coatings or when the nozzle is fabricated with ceramic airfoils and endwalls.
In some turbine nozzle manufacturing methods, each of the vanes and endwalls is separately manufactured and, therefore, subject to inherent manufacturing tolerances. These tolerances are additive and “stack-up” during assembly of the nozzle, which can result in throat area variation. Variations in throat area between adjacent vanes can provide undesirable aero-mechanical excitation pressure forces which may lead to undesirable vibration of the rotor blades disposed downstream from the nozzle. This in turn can lead to engine performance and life reductions.
As can be seen, there is a need for improved methods of modifying nozzle area. Further, methods are needed wherein hardware expense can be reduced while optimal vector diagram to the blade can be maintained. Additionally, methods of reducing variation in throat area are needed.